In gas turbine engines the performance of the basic engine cycle, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbine at the highest possible temperature. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature will always produce more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the lifespan of an uncooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling for many of its components.
In modern gas turbine engines the high pressure (HP) turbine gas temperatures are now much hotter than the melting point of the turbine blade materials commonly used, which therefore necessitates efficient cooling of the HP turbine components. In some engine designs the intermediate pressure (IP) and low pressure (LP) turbines are also cooled. During its passage through the turbine the mean temperature of the gas stream decreases as power is extracted. Therefore the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the HP stage(s) through the IP and LP stages towards the exit nozzle.
Internal convection and external films are the primary methods of cooling the gas-path components, for example aerofoils, platforms, shrouds and shroud segments. HP turbine nozzle guide vanes (NGVs) in particular consume the greatest amount of cooling air in high temperature engines. Typical NGVs comprise forward and rear cavities for passage of cooling air therethrough. HP turbine blades typically use about half of the NGV flow, whereas the IP and LP stages downstream of the HP turbine use progressively less cooling air.
FIG. 1 of the accompanying drawings is an isometric cut-away view of a typical single-stage cooled gas turbine engine 1, showing the NGVs 2 (with their respective aerofoils 3), turbine rotor blades 4 (with their respective aerofoils 5), inner and outer platforms 6, 8, HP turbine disc 10, and pre-swirl nozzles 12, as well as the cover-plates and lock plates arrangements including HP turbine support casing 14 and shroud segments 16.
The HPT blades 4 and NGVs 2 are cooled by using high pressure (HP) air from the compressor that has by-passed the combustor and is therefore relatively cool compared with the gas temperature. Typical cooling air temperatures are in the range of from about 800 to about 1000 K. Gas temperatures can be in excess of about 2100 K.
The cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Extracting coolant flow therefore has an adverse effect on the engine operating efficiency. It is thus important to use this cooling air as effectively as possible.
In modern engines the ever increasing gas temperature level combined with a drive towards flatter combustion radial profiles (in the interests of reduced combustor emissions) have resulted in an increase in local gas temperatures experienced by the NGV aerofoils and their platforms. However, the increased coolant flow levels required to accommodate these changes in local gas temperature have a detrimental effect on the local feed pressure supplying the internal cooling holes of the NGV aerofoils 3. In order to maintain a safe backflow pressure margin, higher pressure drop levels across the combustor are becoming a necessity.
The last ˜10 years has seen a significant rise in the inlet gas temperature and overall engine pressure ratio in newer engine designs, and this has brought a new raft of problems in the turbine section of the engine. The turbine life is typically limited by the lives of the hot section aerofoil components. The HPT NGVs are subject to the hottest gas temperatures and thus consume the highest quantity of cooling flow in order to ensure mechanical integrity.
A leading edge (L/E) region of each NGV aerofoil is also subject to the highest local levels of external heat transfer coefficients, and therefore the L/E region requires special treatment in terms of convective cooling levels. The most commonly employed arrangement involves the provision of plural rows of very steeply inclined cooling holes, located very close together, such as at 20 in FIG. 1. This arrangement is known as a “showerhead”, and its operation is known as “showerhead cooling”.
The coolant air flow levels passing through these showerhead holes dictate the rate of heat removal from the L/E region. However the pressure ratio across these cooling holes is not very high, and therefore under some engine conditions there is a danger of hot gas ingestion into at least the forward cooling cavity of the NGV. If this occurs then the consequences can be disastrous, owing to the resulting failure to provide the necessary cooling. The most likely location where this situation may occur is at the entrance to the forward NGV cooling chamber that feeds the showerhead cooling holes. This is because the flow level at the entrance is at a maximum value compared with the flow area of the feed passage. Consequently the flow velocity, or Mach number, of the coolant air flow is at its highest value in this region and thus the corresponding static feed pressure is at its lowest value. Hence the local pressure ratio across the showerhead cooling holes is also at its lowest level. In order to guarantee that no hot gas is ingested, the “backflow margin” is generally set to a value in the range of about 1.5 to about 2.0%, depending on the confidence levels associated with the pressure data and the familiarity of the cooling geometry. The backflow margin is defined as:backflow margin=(Pstatic feed−P40)/P40×100%,where Pstatic feed is the pressure of the coolant air flow entering the NGV forward cooling chamber and P40 is the pressure of the hot gas flow exiting the combustor.
The cooling holes located on either side of an aerofoil stagnation point are most at risk. The stagnation point tends to move around due to unsteadiness of the flow, and therefore a stagnation region is created where the local static feed pressure of the coolant air flow is equal to the total gas flow pressure Pt40.
In order to ensure that the static feed pressure never falls too low, the leading edge feed passage is generally divided into two separate feed chambers and these have separate coolant air supplies: one fed from the outboard side and the other fed from the inboard side. In addition the respective entrances to these feed passages are generally shaped like the mouth of a bell (i.e. “bell-mouthed”) in order to keep the entrance losses to a minimum.
FIGS. 2(a) and 2(b) of the accompanying drawings show one example of the above typical known arrangement of HPT NGV aerofoils 3, internal cooling geometry and coolant feed systems. FIG. 2(a) is a cross-sectional view of a typical HPT NGV aerofoil cooling scheme with forward 3F and rearward 3R cooling chambers. The forward chamber 3F is a dual feed system, i.e. fed from both outboard and inboard sources, whereas the rear chamber 3R is fed from the inboard source only. FIG. 2(b) is an isometric cut-away view of the NGV aerofoil segment showing the internal cooling scheme features and coolant flows, which are as follows: 30 represents the forward chamber outboard coolant air feed to the leading-edge showerhead holes from which the coolant air exits as at 20; 32 represents the forward chamber inboard coolant air feed; 34 represents the rear chamber inboard (only) coolant air feed; 36 represents an impingement plate having holes therein through which coolant air passes from the rearward chamber 3R to cool the rear side section of the NGV aerofoil; 38 represents cooling air exiting the rear trailing edge of the aerofoil via slots in a pedestal bank, and 39 represents film cooling of exited air from the showerhead holes across the exterior surface of the aerofoil 3. Also shown is a sheet metal baffle plate 35 within the forward chamber 3F for preventing the inboard or outboard coolant sources from dominating the coolant feed system.
This design shown in FIGS. 2(a) and (b) is an example of the simplest of forward cooling chambers 3F, with no heat transfer augmentation features, and fed from both inboard and outboard coolant sources. Also for simplicity the entrances to the forward 3F and rearward 3R chambers are shown here as sharp-edged, with no “bell-mouth” shape.
FIG. 3(a) of the accompanying drawings is a cross-sectional view through another example of a known HP turbine NGV aerofoil cooling scheme, where the NGV aerofoil 3 again comprises forward 3F and rearward 3R cooling chambers. However these chambers differ from those shown in FIGS. 2(a) and 2(b) in that they have mounted therein respective sheet metal impingement tubes 43F, 43R, which are inserted therein from one end (outboard or inboard) and welded in place. The purpose of the impingement tubes 43F, 43R is to provide a plenum from which the coolant air is bled through a series of holes 48 as “impingement jets”, generally arranged in rows, which impinge cooling air onto the inner surface of the respective cast chamber 3F, 3R. These impingement tube devices 43F, 43R enable the designer to target specific locations within the chambers 3F, 3R that correspond to the external heat load, thereby enabling specific localised cooling requirements to be optimised. The coolant air is then bled out onto the exterior surface of the aerofoil through film cooling holes machined into the casting walls, to provide a thermal barrier of cool air which insulates the metal of the aerofoil from the hot gas from the combustor. FIG. 3(a) shows the coolant air being fed into the sheet metal impingement tube inserts 43F, 43R from both inboard and outboard sources 30, 32, 34. Again, there may be provided a sheet metal baffle plate (not shown) located inside at least the forward impingement tube insert 43F for preventing the inboard or outboard coolant sources from dominating the feed system.
FIG. 3(b) of the accompanying drawings is a cross-sectional view through another example of a known HP turbine NGV aerofoil cooling scheme, again with forward and rearward cooling chambers 3F, 3R. However, in this case the chambers 3F, 3R do not contain sheet metal impingement tube inserts as in FIG. 3(a). Instead, in this design impingement cooling air is bled through impingement jets 58 in an additional internal cast wall 53F, 53R located in close proximity to the suction side walls 50 of the aerofoil. FIG. 3(b) shows the coolant air again being fed into the forward cooling chamber 3F from both inboard 32 and outboard 30 sources, while the rear chamber 3R is typically fed from one end only, usually the inboard end 34, where the cooling air is cleaner.
These known designs of NGV aerofoil cooling arrangements all suffer from various problems, shortcomings or limitations. For example:                Higher gas temperatures experienced by the NGV aerofoils result in higher coolant flow requirements. As recent modern engine designs have been developing, aerofoil shapes and sizes have not been changing in proportion to the increased coolant flow levels that are required. As a result, the local velocity of the coolant in the outboard and/or inboard feed chambers needs to increase, in particular that/those supplying the forward NGV cooling chamber that supplies the L/E showerhead holes, where local pressure levels are critical.        
Dual-end feed forward cooling chambers have proved beneficial to some extent in keeping the inlet velocity (Mach no.) down to relatively low levels, but these improvements have already been offset by the increased flow demands, so any further improvement based on such features is limited.                “Bell-mouthed” entrance shapes have also played a useful part in keeping the local inlet velocity (Mach no.) at a low level by eliminating or reducing inlet separation, but again this feature only gives limited improvements.        According to current knowledge in the art, the only ways that remain for the designer to ensure that the cooling chamber feed pressure is kept at a safe level above that in the gas-path is to accommodate the increased flow requirements by designing a “fatter” aerofoil shape at the root or tip sections thereof, in order to increase the inlet flow area of the forward cooling chamber, or to increase the pressure drop across the combustor (P30-P40). However, both such design changes would seriously affect the efficiency of the turbine or cycle of the engine.        